The present invention relates to attitude control systems employed in space vehicles.
Space vehicles are generally provided with attitude control systems for maintaining the vehicle in a desired attitude with respect to some set of reference axes which do not rotate with the vehicle. In most cases, the reference axes either are aligned with the vehicle orbit or fixed with respect to the sun or stars.
Conventional attitude control systems contain a plurality of channels, there generally being one channel associated with each vehicle body axis, and each channel may contain an electromechanical actuator which can be in the form of a reaction wheel assembly, a momentum wheel assembly, or a control moment gyro assembly. Each type of assembly includes a wheel which is rotated at high speed to produce a force tending to oppose disturbance forces acting on the vehicle.
During prolonged operation, such actuators tend to become saturated with unwanted disturbance energy.
In the case of reaction wheel assemblies, saturation occurs when the wheel drive motor reaches its limit speed.
Momentum wheel assemblies become saturated when the momentum wheel and the space vehicle precess together under the influence of a disturbance torque until the resulting angular displacement of the vehicle from a desired orientation becomes unacceptably large, or until the speed of rotation of the momentum wheel deviates from the desired, fixed speed by a predetermined amount.
In the case of control moment gyros, saturation occurs when the wheel gimbal displacement becomes such that required control torque outputs can no longer be generated, or the gimbals reach the mechanical stops that limit their pivoting movement. Such gyro devices can have either one or two degrees of freedom. When the device has two degrees of freedom, the actuator includes two gimbal pivoting motors and two loops, or channels, each associated with a respective degree of freedom.
To desaturate a reaction wheel assembly, the usual practice is to apply a voltage to the associated drive motor amplifier in a manner to reduce the speed of the reaction wheel. In the case of a momentum wheel assembly, a voltage is applied to the momentum wheel drive motor amplifier in order to return the momentum wheel speed to its nominal value. Moreover, torque is applied to the vehicle by means other than the momentum wheel assembly in order to return the vehicle and the momentum wheel to the desired attitude. For a control moment gyro assembly, desaturation is effected by acting on the or each gimbal motor to drive the associated gimbal toward its null position.
Whenever desaturation is being effected by applying voltage to a drive motor amplifier, the result is to impose a disturbance torque on the space vehicle. Therefore, it is usually necessary to provide an additional source of stabilizing torque on the vehicle, at least during the desaturation phase.
Such an additional stabilizing source may be of the passive type, such as gravity gradient booms or aerodynamic fins, or active devices not of the electromechanical variety, such as gas jet thrusters or magnetic torque rods.
It is the general practice in the art to effect momentum management, or energy desaturation, by deriving signals representative of wheel motion. In the case of a reaction wheel, only a signal representative of wheel speed is needed. If the actuator employs a momentum wheel, multiple signals are needed: signals representative of the angular attitude of the satellite; and a signal representative of wheel speed. If the actuator employs a control moment gyro, the signal is representative of the angular position of a respective gimbal. The wheel motion signal is applied to a computer in which a complicated algorithm is implemented to produce a torque command voltage that will drive the wheel, or the gimbal, in the manner required to achieve desaturation.
All facets of space vehicle attitude control are described in a text edited by James R. Wertz, Spacecraft Attitude Determination and Control, D. Reidel Publishing Company, Dordrecht, Holland, 1978.
Various types of attitude control systems are described in the following publications:
V. N. Branets, et al., Development Experience of the Attitude Control System Using Control Moment Gyros for Long-Term Orbiting Space Stations, 38th Congress of The International Astronautical Federation, Brighton, United Kingdom, Oct. 10-17, 1987, pp 1-8;
H. F. Kennel, Steering Law for Parallel Mounted Double-Gimbaled Control Moment Gyros--Revision A, NASA Technical Memorandum, TM-82390, January 1981, available from NTIS.
Other publications describing attitude control system and energy desaturation, or momentum management, techniques are:
Bendix Research Laboratories, Southfield, Mich., Technical Report BRL/TR-73-6768, CMG/TACS Control System Hybrid Simulation, September 1973;
John R. Glaese, et al., Torque Equilibrium Attitude Control for Skylab Reentry, NASA TM-78252, November 1979, available from NTIS;
Robert O. Hughes, Conceptual Design of Pointing Control Systems for Space Station Gimballed Payloads, presented as Paper 86-1986 at the AIAA Guidance, Navigation, and Control Conference, Williamsburg, Va. Aug. 18-20, 1986, pp 78-87;
Henry H. Woo, et al., Momentum Management Concepts for a Space Station, presented as Paper 86-2047 at the AIAA Guidance, Navigation, and Control Conference, Williamsburg, Va., Aug. 18-20, 1986, pp 277-286; a revised version was published in J. Guidance, Vol. 11, No. 1, January--February 1988, AIAA, pp 19-25; and
Henry N. Woo, et al., Preliminary Evaluation of a Reaction Control System for a Space Station, presented as Paper 86-2152 at the AIAA Guidance, Navigation, and Control Conference, Williamsburg, Va., Aug. 18-20, 1986, pp 538-546.
The use of gravity gradient booms for spacecraft stabilization is described in the following publications:
D. K. Anand, et al., Attitude Performance of Some Passively Stabilized Satellites, Journal of the British Interplanetary Society, Vol. 26, 1973, pp 641-661;
David L. Blanchard, Flight Results from the Gravity-Gradient-Controlled RAD-1 Satellite, presented as Paper 86-2140 at the AIAA Guidance, Navigation, and Control Conference, Williamsburg, Va., Aug. 18-20, 1986, pp 479-487; and
R. V. Davis, et al., Flight Experience and Application of Earth-Orbiting Gravity Gradient Stabilization Systems. (Proceedings of the Sixteenth International Astronautical Congress, Anthens, Greece, International Astronautical Federation, Vol. II, 1966), pp 293-300 and four pages of Figures.
FIG. 1 illustrates one known type of attitude control system in which spacecraft attitude is maintained by at least one auxiliary passive torque source during actuator desaturation. In the drawing, mechanical couplings are illustrated by solid lines and electrical couplings by broken lines, and one attitude control channel is illustrated. Spacecraft frame 2 is mechanically coupled to an electromechanical actuator 4 and an auxiliary passive torque source 6. In normal operation, actuator 4 will apply a torque tending to oppose deviations in the attitude of frame 2 from its desired orientation relative to at least one reference coordinate, or axis. This is indicated by the negative sign in the line coupling actuator 4 to frame 2. Thus, in normal operation, actuator 4 tends to stabilize frame 2.
Auxiliary passive torque source 6 may, depending on the manner in which the satellite attitude is to be controlled, apply either a stabilizing or destabilizing torque to frame 2 during various orbital phases.
The true attitude angle and true angle rate of spacecraft frame 2 are sensed by angle motion sensors 8 which produce output signals indicative of those parameters. Those signals are applied to a computer 10 which derives, on the basis of those values, for a given attitude control channel, a torque command voltage which is applied to a power amplifier 14 in actuator 4. The output voltage produced by amplifier 14 is supplied to an electric motor 16 to cause that motor to generate a torque which acts in a direction to return frame 2 to the desired attitude about the coordinate axis associated with the attitude control channel.
The rotor of motor 16 is coupled to a momentum storage element 18 which rotates together therewith. The motion of element 18 is indicative of the disturbance energy stored therein and is sensed by a sensor 20 which supplies a signal indicative of measured motion to computer 10.
If actuator 4 is a reaction wheel assembly or a momentum wheel assembly, element 18 is the wheel, which is driven by motor 16, and sensor 20 produces a signal representative of the direction and speed of wheel rotation. If actuator 4 is a control moment gyro assembly, element 18 is a pivotally mounted gimbal supporting the gyro rotor and pivoted by motor 16, and sensor 20 produces a signal representative of the direction and magnitude of the angular deviation of the gimbal from its null position.
Computer 10 processes the information provided by the measured motion signal according to a relatively complicated algorithm in order to apply to power amplifier 14 a signal which drives motor 16 in a manner to dissipate the disturbance energy stored in element 18.
A system of the type illustrated in FIG. 1 would be employed, for example, in an earth-oriented satellite, where the gradient of the earth's gravity is the available source of auxiliary passive restoring torque with respect to the pitch and roll attitude control channels, while the inertial reaction of the spacecraft frame to the rotation of the satellite around the earth is the available source of auxiliary passive restoring torque in the yaw channel.
In the case of a satellite which is orbiting the earth but which is oriented in a reference frame fixed in inertial space, for example a sun-oriented or star-oriented satellite, the earth's gravity gradient is the available source of auxiliary passive restoring torque for all attitude control channels. However, in this case, the ability of the auxiliary passive source of restoring torque to impose a stabilizing influence on the spacecraft frame is intermittent since, as the satellite orbits the earth, the gravity gradient will alternately generate restoring torque and disturbance torque. In this case, electromechanical actuator desaturation must be interrupted during the time periods when the gravity gradient is generating a disturbance torque because actuator desaturation at those times would destabilize the satellite attitude.
FIG. 2 illustrates a second known type of attitude control system which is identical to that shown in FIG. 1 except for the type of auxiliary actuators employed. In the case of the arrangement shown in FIG. 2, the auxiliary actuators 22 are not passive torque sources, but instead are actuators of the active type, such as gas jet thrusters or magnetic torque rods. These actuators 22 would be controlled by appropriate signals from computer 10. In the arrangement shown in FIG. 2, desaturation of the electromechanical actuators 4 would be effected in the same manner as described above in connection with FIG. 1.
The desaturation techniques described above have a number of disadvantages which are connected with the fact that they must be controlled by a digital computer operating according to a relatively complicated algorithm. Specifically, development, testing and correction of the algorithm is a relatively time consuming, and hence expensive, procedure, particularly since they must allow for programmed attitude maneuvers and must take account of the fact that the angle rates at which spacecraft rotate around their own centers of mass, in response to torques generated by the actuators, are finite. Moreover, when a new spacecraft is being developed, its performance parameters in space can only be estimated. Once the spacecraft has been placed into orbit, adjustments to the algorithm employed to control desaturation, due to differences between estimated and actual performance, are difficult, if not impossible, to effectuate.
Furthermore, the existing techniques raise the possibility of overshoot in the torque commands to the electric motors driving the momentum storage elements in the actuators which are being desaturated, resulting in unwanted energy being put back into the elements.